FIG. 1 shows a gas turbine engine as is known from the prior art. With reference to FIG. 1, a gas turbine engine is generally indicated at 100, having a principal and rotational axis 11. The engine 100 comprises, in axial flow series, an air intake 12, a propulsive fan 13, a high-pressure compressor 14, combustion equipment 15, a high-pressure turbine 16, a low-pressure turbine 17 and an exhaust nozzle 18. A nacelle 20 generally surrounds the engine 10 and defines the intake 12.
The gas turbine engine 100 works in the conventional manner so that air entering the intake 12 is accelerated by the fan 13 to produce two air flows: a first air flow into the high-pressure compressor 14 and a second air flow which passes through a bypass duct 21 to provide propulsive thrust. The high-pressure compressor 14 compresses the air flow directed into it before delivering that air to the combustion equipment 15.
In the combustion equipment 15 the air flow is mixed with fuel and the mixture combusted. The resultant hot combustion products then expand through, and thereby drive the high and low-pressure turbines 16, 17 before being exhausted through the nozzle 18 to provide additional propulsive thrust. The high 16 and low 17 pressure turbines drive respectively the high pressure compressor 14 and the fan 13, each by suitable interconnecting shaft.
It is known that turbine engine efficiency is closely related to operational temperatures and acceptable operational temperatures are dictated to a significant extent by the material properties of the components. With appropriate cooling it is possible to operate these components near to and occasionally exceeding the melting points for the materials from which they are constructed in order to maximise operational efficiency.
Generally, coolant air is taken from the compressor stages of a gas turbine engine. This drainage of compressed air reduces the quantity available for combustion and consequently, engine efficiency. It is desirable to use coolant air flows as effectively as possible in order to minimise the necessary coolant flow to achieve a desired level of component cooling for operational performance. Intricate coolant passageways are provided within engine components and are arranged to provide cooling. The coolant passes through these passageways and is typically delivered to cavities in regions requiring cooling. Delivery into a cavity is often by nozzle projection which serves to create turbulence with hot gas flows for a diluted cooling effect.
One area where compressed coolant air is known to be used is between stages in a gas turbine engine. The coolant air is typically delivered into a cavity between discs of adjacent turbine stages. The discs may be rotor discs. The cavity may be positioned radially inwardly of a stationary nozzle guide vane which is arranged axially (i.e along the engine axis) between the discs. The coolant may be swirled to complement the direction and speed of rotation of a rotor disc on delivery to the disc surface.
A prior art arrangement is shown in FIG. 2 which is a schematic cross-section of a prior cooling arrangement for a turbine inter-stage. As shown, first blade 1 forms a shank with a locking plate 2 presented across the root 3 of the blade 1. Seals 4 are provided in the form of a labyrinth seal arrangement with coolant airflow (compressed air which has bypassed the combustor) in the direction of arrowhead 5. The coolant air travels radially outwardly (upwardly in the view shown) and into the cavity 6 formed between the mounting disc 7 for the blade 1 and the bottom of a nozzle guide vane dividing the axially adjacent turbine stages. As can be seen there is a gap 8 through which hot gas is ingested into the cavity 6. The coolant air 5 has been arranged to prevent excessive hot gas ingestion, the direction of which is represented by arrowhead 9. This can be achieved by appropriate balancing of pressures between the hot gas and coolant in the region. The locking plate 2 acts to secure location of the blade shank 1 such that coolant flow 5 is contained or at least restricted below the blade shank 1. An area 10 adjacent the lock plate 2 allows coolant air to flow across it at its surface to provide cooling. The lock plate 2 is segmented, the gaps between the segments allowing coolant leakage into the cavity 6. It will be understood that unwanted hot gas ingestion occurs when the coolant flow supplied to the rim gap is less than the critical value required to seal the rim gap. In the case of an inter-stage seal cavity where the labyrinth seal clearance is such that the cooling flow is drawn off to the lower pressure “sink”, downstream of the stage nozzle guide vane, leaving the gap at the rear of the upstream rotor short of the necessary flow requirements to create the seal at the annulus. Thus, as engines complete more and more service cycles and the inter-stage seals tend to wear there is also an increase in the clearances and redistributing the normally fixed level of coolant flow towards the rear stator well. This increases the risk of hot gas ingestion in the front of the well. Thus, pressure differentials between the coolant flow and hot gas need to be carefully controlled if engine efficiency is to be optimised.
There is a balance between the cooling supply and hot gas ingestion dependent upon many factors including the static pressure in the gas turbine annulus, the losses in the cooling air feed system, any flow dependent on a vortex, rotating hole, clearance diameters or seal clearance subject to a combination of rotor speeds, the main annulus pressure ratios and transient effects such as seal clearances. In such circumstances, a range of conditions over which hot gas ingestion may occur and the level of ingestion will vary.
With ever increasing engine size and higher operating temperatures and engine speeds, pressure losses in the air system increase and coolant flows become less effective and more difficult to control. There is a desire to further improve efficiency of flow of cooling air.